De-icing apparatus for compressors



N. BURGESS DE-ICING APPARATUS FOR COMPRESSORS Filed Dec. 20 1949Inventor: NeH Burgess,

His Attorney.

M Z 3 9 l 1/ m a Z I n u Sept. 20, 1955 United States Patent ()fiice 2,718,350 Patented Sept. 20, 1955 DE-ICING APPARATUS FOR COMPRESSORS NeilBurgess, Melrose, Mass., assignor to General Electric Company, acorporation of New York Application December 20, 1949, Serial No.134,024

7 Claims. (Cl. 230-122) This invention relates to gas turbine powerplants, and more particularly to means for preventing ice accumulationin the inlet passage of such power plants.

A gas turbine power plant for the propulsion of aircraft may comprise anaxial flow compressor and combustion apparatus utilizing compressed airsupplied thereto by the compressor to provide gases under pressure andat elevated temperature for driving the turbine-compressor combination,and a nozzle utilizing gases leaving the turbine to form a propulsionjet. The details of such a power plant are described with greaterparticularity in a copending application of Alan Howard, Serial No.541,565, filed June 22, 1944, now Patent No. 2,711,074, dated June 21,1955, and assigned to the assignee of the present application.

in the operation of such a power plant, it is desirable in the interestof good efliciency to provide a plurality of guide vanes in the inletpassage to direct the entering air to the compressor rotor at thecorrect angle. If the power plant is operated under adverse atmosphericconditions there may be considerable tendency for ice to form andaccumulate on the guide vanes. If any appreciable quantity of iceaccumulates on the guide vanes, the compressor inlet cross-sectionalarea is reduced, thus reducing the quantity of air flowing through thepower plant. A reduction in the air flow will decrease the thrust orpower output of the power plant, and since a reduced quantity of air isflowing through the combustion apparatus, the temperature in thecombustion, turbine, and exhaust sections of the power plant mayappreciably exceed their normal operating values with resultant damageto the power plant.

A second problem in connection with the operation of power plants of thetype described is in the somewhat unrelated problem of balancing axialforces on the compressor rotor in order to minimize bearing loads. Inpower plants of this type it is often desirable to provide at least onethrust balancing piston as part of the compressor rotor structure and todirect fluid pressure against the piston to obtain a balancing force inopposition to said axial forces. Particularly in aircraft service whereweight considerations are of the utmost importance, it is extremelydesirable to construct the wheels or blade carrying discs at the lowpressure end of the compressor rotor of a light weight, non-ferrousmaterial, such as magnesium or aluminum. For the same reason it isdesirable to obtain the pressurized fluid for the force balancing systemfrom the compressor itself, and thus it is desirable to connect thebalancing piston to a point of suitable pressure in the compressor flowsystem. This practice is subject to the disadvantage, however, thatnon-ferrous metals of the type described have relatively littlestructural strength at temperatures exceeding 200 Fahrenheit, and sincethe temperature of the compressed fluid may easily exceed this valuefollowing compression, the rotor structure must be made heavier thanwould otherwise be necessary, or means for cooling the pressurized fluidsupplied to the balance piston must be provided.

Accordingly, it is an object of the invention to provide improved meansfor preventing ice formation at the compressor inlet and at the sametime for cooling fluid supplied to the compressor balance piston.

Another object of the invention is to provide improved means forde-icing the compressor inlet guide vanes of an aircraft power plant.

Still another object is in the provision of means for preventing iceaccumulation at the inlet of an aircraft power plant and for supplyingcooled fluid for balancing axial forces acting on the rotor in such apower plant without incurring losses in the performance thereof.

Another object is in the provision of improved means for balancing axialforces on the rotor of an aircraft power plant without material increasein the weight of the power plant and without requiring special heatexchange apparatus.

Other objects and advantages will be apparent from the followingdescription taken in connection with the accompanying drawings, in whichFig. l is a diagrammatic view, partly in section, of a compressorportion of an aircraft power plant illustrating one embodiment of theinvention; Fig. 2 is an enlarged sectional view of a portion of thecompressor inlet passage shown in Fig. l; and Fig. 3 is a partialsectional view, taken on line 3-3 of Fig. 2 looking in the directionindicated by the arrows, showing the inlet guide vanes and a supportstrut.

Referring now to Figs. 13, a compressor section of a gas turbine powerplant is indicated generally at 1. As previously indicated, a gasturbine power plant for the propulsion of aircraft may also include acombustion section, a turbine section, and a propulsion nozzle. However,none of these components are essential to an understanding of thepresent invention and therefore are not shown. The compressor section 1comprises a stator 2, a rotor 3, and an inlet 4. A plurality of axiallyspaced rows of stationary blades 5 are secured to the stator, asindicated in the drawing. Rotor 3 comprises a shaft 6 on which aresecured a plurality of wheel or disc members. Each of discs 7 is incontact with an adjacent disc at the periphery thereof to form asubstantially continuous surface 8. An inner surface 9 of stator 2 andsurface 8 are coaxial and in spaced relation to define an annular flowpassage through the compressor. Separate rows of moving blades 10 aresecured to each of the peripheral portions of discs 7 and, as indicatedin the drawing, the separate rows of moving blades 10 are disposedbetween adjacent rows of stator blades 5. Axially extending walls 11, 12are secured to the stator 2 and are supported in spaced relation by aplurality of struts 13 to form an annular inlet passage connecting inlet4 and the compressor flow passage defined by rotor surface 8 and statorsurface 9. A plurality of inlet guide vanes 14 are provided in the inletpassage in order to direct fluid at the proper angle to the first row ofmoving blades 10, as indicated in Fig. 3.

In order to support vanes 14, inner and outer shroud bands 15, 16 areprovided. The inner and outer bands are punched to receive the inner andouter end portions of vanes 14 which project entirely through the bands.Since ice tends to accumulate on the guide vanes under adverse weatherconditions, it is desirable to provide means for heating the vanes toprevent such accumulation. This is accomplished in accordance With theinvention by providing hollow guide vanes having a passage 17 extendingtherethrough, as indicated in Fig. 3 and by the broken lines 18, 19 inFigs. 1 and 2. The hollow vanes are secured to bands 15, 16 by weldingor in any other desired manner which will not obstruct the flow of fluidthrough passages 17. The outer end portions of vanes 14 are connected inparallel flow relation by the provision of a U-shaped member 20 securedto band 16 to provide a plurality of radially extending passages asindicated in the drawings. Recessed portions 21, 22 are provided inwalls 11, 12, respectively, for locating and securing the vane and bandassembly in spaced relation with respect to the first row of movingblades 10.

In order to provide a supply of heated fluid under pressure to passages17, a conduit 23 connects the vane assembly to the compressor flow pathat any suitable location 23a at which the desired fluid pressure levelis obtained. Conduit 23 registers with an opening 24 in U-shaped memberto provide a continuous flow path from the compressor passage throughconduit 23 and passages 17 in the guide vanes.

The rotor shaft 6 is rotatably supported in suitable bearings 25, one ofwhich is shown. In order to minimize thrust load on the bearingsresulting from unbalanced forces acting on the compressor rotor in anaxial direction, a balance piston or face 7a against which fluidpressure can be directed is formed by the provision of an axiallyextending peripheral portion 26 of the first rotor disc 7. An axiallyand radially extending wall portion 27 is secured to wall 11 to form asupport for bearing Y and, in cooperation with band 15, wall portion 27also assists in forming a pressure chamber 28 containing the balancepiston 7a defined by peripheral portion 26. Sealing means 29, 30 whichmay be of any wellknown type are provided to prevent excessive leakageof pressurized fluid through the clearance space between wall portion 27and shaft 6, and to control the rate at which fluid flows between band15 and the rotating peripheral portion 26, respectively.

For reasons which will become apparent as the description proceeds, itmay be desirable under particularly adverse icing conditions to providemeans for connecting pressure chamber 28 to the atmosphere withoutexcessive reduction of the chamber pressure. As illustraded in Figs. 1and 3, strut 13 and wall 11 are provided with a flow passageway 31therein. An opening 32, which registers with passageway 31, is providedin wall portion 27 for establishing communication between chamber .28and passageway 31. A valve 33 for controlling the fiow of air throughpassageway 31 is series connected thereto by conduit 34. Valve 33 isnormally closed to prevent any flow through passageway 31.

In operation air is drawn from the atmosphere through annular inletpassage 4 and is compressed by the compressor section 1. The combustionsection (not shown) utilizes this compressed air to furnish hightemperature motive fluid for the turbine (not shown) which drives thecompressor. Air under pressure is conveyed from the compressor to theinterior passages 17 of guide vanes 14 through conduit 23. The airpassing through the compressor is heated considerably by the compressionprocess and this flow of heated air through the hollow portions of vanes14 warms the vanes and prevents the accumulation of ice thereon. Sincethe air enters the compressor through inlet passage 4 at relatively highvelocity and at a substantially lower temperature than that of the fluidflowing through passages 17, the combination of these several factorscauses hollow vanes 14 to function as an effective heat transferapparatus, especially if vanes 14 are provided with thin walls. Becauseof the transfer of heat to the entering air from the heated vanes 14,the air discharged from passages 17 into pressure chamber 28 is cooledto a temperature substantially below the temperature of the air inconduit 23 and thus the first rotor disc 7 is protected againstexcessive temperature. Normally, the rate at which air is supplied topressure chamber 28 through conduit 23 and hollow vanes 14 for thrustbalancing purposes is great enough to cause suflicient heating of vanes14 to prevent ice accumulation thereon. Valve 33 is therefore allowed toremain in its normally closed position, and the compressor is protectedagainst icing without the use .of additional air for heating purposesover the air normally supplied to the balance piston-7a. Thus, theinvention provides adequate de-icing protection to the power plant andat the same time provides effective cooling of the balance piston airfor the protection of the first rotor disc without the use of additionalheat exchange apparatus and the attendant increase in weight of thepower plant and without loss in performance thereof.

Under unusually severe operating conditions, the normal rate of flow ofheated air for thrust balancing purposes through the hollow portions ofvanes 14 may be inadequate to provide the desired degree of de-icingprotection. In such instances the temperature of vanes 14 and hence theeffectiveness of the ice prevention apparatus can be increased byopening valve 33, thereby increasing the flow of heated air through thevanes.

Thus, it will be seen that the invention provides antiicing protectionfor an aircraft power plant, and at the same time cools pressurized airwhich is supplied to a compressor balance piston thereby minimizingtemperature differentials in the region of the first stage rotor discwithout loss in performance of the power plant; and the strength andreliability of certain critical parts are thereby substantiallyincreased without the use of additional heat exchange apparatus.

While a particular embodiment of the invention has been illustrated anddescribed, it will be obvious to those skilled in the art that variouschanges and modifications may be made without departing from theinvention, and it is intended to cover in the appended claims all suchchanges and modifications that come within the true spirit and scope ofthe invention.

What I claim as new and desire to seecure by Letters Patent of theUnited States is:

1. In a gas turbine power plant including a compressor having a bladedrotor, means for preventing ice accumula tion at the compressor inletcomprising first walls surrounding an unbladed portion of the rotor anddefining therewith a substantially closed chamber, other wallssurrounding said chamber defining an annular inlet for conveying fluidto the compressor, a balance piston within said chamber and connected tothe rotor, a plurality of hollow vanes extending across said inletgenerally in a direction normal to said walls and with the hollowportions of said vanes in communication with said chamber, means forsupplying fluid under pressure to said piston including conduit meansconnected to said hollow portions and to the compressor, and sealingmeans between said first walls and said rotor for preventing excessiveleakage of fluid from said chamber into said annular inlet.

2. In an aircraft power plant, multi-stage air compressor including astator carrying a plurality of spaced rows of stationary blades and arotor coaxial therewith and carrying a plurality of rows of movingblades disposed between adjacent rows of stationary blades, said rotorand stator cooperatively arranged to form a flow passage through thecompressor, said rotor having an unbladed peripheral portion adjacentthe first row of moving blades, first walls substantially enclosing saidperipheral portion, other walls defining an annular inlet passage incommunication with said flow passage and surrounding said first walls, aplurality of hollow vanes projecting through said first walls andextending across said inlet passage in a direction substantially normalto the direction of flow therein, means for supplying heated air underpressure including conduit means connected to said hollow vanes and tosaid compressor flow passage, and sealing means between said first wallsand the rotor for preventing excessive air leakage from said chamber.

3. In an aircraft power plant, a compressor having a stator carrying aplurality of axially spaced rows of blades and having a rotor coaxialwith said stator and carrying a plurality of spaced rows of bladesdisposed between adjacent rows of stationary blades, said rotor and saidstator being cooperatively disposed to form a compressor flow passage, abalance piston secured to the rotor at the upstream side of the firstrow of rotor blades and rotating coaxially therewith, first wallssurrounding said balance piston and forming a substantially closedchamber containing said piston, walls defining an annular inlet passagesurrounding said chamber, a plurality of hollow blades extendingsubstantially radially across said inlet and having end portions thereofprojecting into said chamber, for supplying fluid under pressure to saidbalance piston including conduit means connecting the other end portionsof said hollow blades to said compressor flow passage at a locationbetween the first and the last row of moving blades, and means forpreventing excessive fluid leakage between said first walls and saidpiston.

4. In a gas turbine power plant including a compressor having a bladedrotor, means for preventing ice accumulation at the compressor inletcomprising first walls surrounding an unbladed portion of the rotor anddefining therewith a substantially closed chamber, other wallssurrounding said chamber defining an annular inlet for conveying fluidto the compressor, a balance piston within said chamber and connected tothe rotor a plurality of hollow vanes extending across said inletgenerally in a direction normal to said walls and with the hollow portions of said vanes in communication with said chamber, means forsupplying fluid under pressure to said piston including conduit meansconnected to said hollow portions and to the compressor, sealing meansbetween said first walls and said rotor for preventing excessive leakageof fluid from said chamber into said inlet, and means for changing therate of fluid flow through the hollow portions of said vanes.

5. In an aircraft power plant, a compressor having a stator carrying aplurality of axially spaced rows of blades and having a rotor coaxialwith said stator and carrying a plurality of spaced rows of bladesdisposed between adjacent rows of stationary blades, said rotor and saidstator being cooperatively disposed to form a compressor flow passage, abalance piston secured to the rotor at the upstream side of the firstrow of rotor blades and rotating coaxially therewith, first wallssurrounding said balance piston and forming a substantially closedchamber containing said piston, walls defining an annular inlet passagesurrounding said chamber, a plurality of hollow blades extendingsubstantially radially across said inlet and having end portions thereofprojecting into said chamber, for supplying fluid under pressure to saidbalance piston including conduit means connecting the other end portionsof said hollow blades to said compressor flow passage at a locationbetween the first and the last row of moving blades, means forpreventing excessive fluid leakage between said first walls and saidpiston, and means for increasing the rate of fluid flow through thehollow blades including conduit means and valve means, said conduitmeans and said valve means being series connected to said chamber forestablishing communication between said chamber and the atmosphere.

6. An axial flow compressor having a casing with at least one row ofradially extending hollow vanes therein, a wall interconnecting theinner end of said guide vanes, a diaphragm extending inwardly from saidwall, a rotor having at least one row of radially extending bladesdownstream of said vanes, and means for discharging compressed air froma point within the compressor downstream of said blades into the saidvanes, in combination with a chamber defined by the wall, the diaphragmand the upstream end of the rotor, and a connection from said hollowvanes to said chamber to supply air from said vanes to said chamber tomaintain a thrust balancing pressure therein.

7. An axial flow compressor having a casing with hollow inlet guidevanes extending substantially radially inwardly therefrom, an annularwall interconnecting the inner end of said guide vanes, a diaphragmextending inwardly from said annular wall, said casing also having rowsof stator vanes extending inwardly therefrom and located downstream ofsaid guide vanes, a rotor having rows of blades alternating with therows of stator vanes, said diaphragm having a bearing therein for theupstream end of said rotor, said rotor having an end wall adjacent tosaid diaphragm, said annular wall, end wall and diaphragm defining achamber therebetween, sealing means between said end wall and diaphragmadjacent the bearing, and other sealing means between said annular walland rotor end wall at a point radially outward of said first sealingmeans to complete the chamber, and means for discharging into said guidevanes a portion of the compressed air from the compressor, the innerends of the guide vanes communicating with said chamber.

References Cited in the file of this patent UNITED STATES PATENTS2,333,053 Stroehlen Oct. 26, 1943 2,429,681 Griflith Oct. 28, 19472,462,600 Boestad et al Feb. 22, 1949 2,469,375 Flagle May 10, 19492,474,068 Sammons et al June 21, 1949 2,474,258 Kroon June 28, 19492,477,798 Grifiith Aug. 2, 1949 FOREIGN PATENTS 619,390 Great BritainMar. 8, 1949

